High pressure ratio axial flow supersonic compressor



March 14, 1961 J. K. KOFFEL EI'AL 2,974,858

HIGH PRESSURE RATIO AXIAL FLOW SUPERSONIC COMPRESSOR Filed Dec. 29, 19552 Sheets-Sheet 1 March 14, 1961 J. K. KOFFEL EI'AL 2,974,858

HIGH PRESSURE RATIO AXIAL FLOW SUPER-SONIC COMPRESSOR Filed Dec. 29,1955 2 Sheets-Sheet 2 VLr V52 Ver Via 15 Z00 Vir -Ver' [HI "En fax-'Jamv ,5. ,(J/OFF'EL C-(AELE-S' A. L/NDLEY United States Patent F HIGHPRESSURE RATIO AXIAL FLOW SUPERSONIC COMPRESSOR John K. Kolfel,Cleveland Heights, Ohio, and Charles A. Lindley, Sierra Madre, Calif.,assignors to Thompson Ramo Wooldridge Inc., a corporation of Ohio FiledDec. 29, 1955, Ser. No. 556,337

7 Claims. (Cl. 230-120) The present invention relates to axial flowfluid compressors and, more particularly, it is concerned with theconstruction of an axial flow compressor capable of producing a highpressure ratio from a single axial flow stage, when operating underrelative supersonic compressor inlet conditions.

With the advent of jet aircraft and other types of aircraft capable ofoperating at or above the speed of sound, the problem of air compressionto provide air for the burning of fuel, increased materially. It wasdiscovered early in testing operations that conventional axial flowcompressors designed for operation in the subsonic range (velocitiesbelow Mach 1) are entirely unsatisfactory for compression of fluidsentering the compressor at velocities at or above Mach 1. Accordingly,test procedures and structures were developed for determining theoptimum compressor structure for use at supersonic velocities. As aresult of this comprehensive design construction and testing programsupersonic axial flow compression of air or the like has been provedpracticable and efficient. In a prior supersonic compressor developed inthis program efficient compression of air entering the system atvelocities above Mach 1 was found practical for providing a pressureratio of approximately 2.5 to 1 for a single stage. While this priorcompressor was found to be very satisfactory relative to flow etficiencyand accordingly was an important advance in the field, the pressureratio for eflicient operation was limited to the above noted value ofapproximately 2.5 for a single stage. While this is an unusually highpressure boost per stage for supersonic compressors theretofore known inthe art, it is very desirable to provide, if possible, a compressorcapable of substantially higher ratio, preferably in the range of 4 to 6in a single compression stage.

Prior to the instant application no such high pressure ratio singlestage compressor of a supersonic type was 'known. In fact, among theleading researchers and writers on the subject substanital doubt existedas to whether or not such a compressor could be constructed. The majordifficulty facing those skilled in the art in obtaining high pressureratios at supersonic speeds was the necessity of providing a high degreeof fluid turn during passage through the rotor efliciently. To ourknowledge, no prior art has successfully overcome the diflicultiesinvolved in turning fluid at supersonic velocities without extreme flowseparation in such a manner as to provide a compressor capable ofproviding extremely high pressure ratios in a single rotor-stator stage.The structure of the present invention satisfactorily provides such acompressor through the use of blading, either in the rotor and stator orin the stator only which is capable of accepting fluid flow above thespeed of sound, decreasing the speed or the velocity of flow downthrough the speed of sound and subsequently turning the fluid flow asubstantial number of degrees subsonically and with diffusion.

' Compressors for supersonic operation having both super- 7 2,974,858Patented Mar. 14, 1961 sonic rotor and supersonic stator, requiring alarge angle of fluid flow turn in the rotor, have previously beenconsidered impractical because of the high turning angle required andthe impossibility, then believed, of providing for such high angles ofturn in the range of supersonic operation. Through experimentation, wehave found that very high turning angles can be incorporated insupersonic blading where such blading provides a normal shock,positioned where the fluid flow passes downwardly through the velocityof sound, with subsequent diflusion, if the subsequent diffusion iscarefully controlled by limiting the expansion of the channel areadownstream of the throat or point of normal shock location to a smallpercentage while the flow is being turned. By further providing highblade solidity in order to provide adequate support for the flow throughthe high turning required and by using sharp blade leading edges havinga small blade included angle in order toprevent the build-up of strongoblique shocks providing, instead, a series of low loss weak obliqueshocks, efiiciency is maintained. Further, the outer and inner passagecontours of both the rotor and the stator are controlled, as willhereinafter more fully be developed, to provide radial outward flow turnto establish radial equilibrium and prevent a substantial pressuregradient, occurring between the inner and outer confines of the fluidflow channel.

By providing blading capable of satisfactorily turning fluid flowthrough large angles of turn and at the same time diffusing through thespeed of sound, it has been possible to provide two types of singlestage compressors capable of producing high pressure ratios heretoforeconsidered impractical by prior art engineers. These comprise first, acompressor of the impulse rotor type having satisfactory overallefliciency has been constructed. The compressor provides a high degreeof flow turn in the compressor rotor without any velocity change andhence without a normal shock in the rotor passageway. In view of thiscondition the fluid leaves the impulse rotor at a velocity not onlysupersonic relative to the stator but supersonic relative to the rotorblade and accordingly substantially higher velocity relative to thestator than in a shock-in-rotor compressor design. The flow leaving theimpulse type rotor must be turned to the axial I direction and reducedin velocity, which functions are achieved through the use of the bladingabove described wherein diffusion is very carefully controlled and thepassageway area expansion is maintained at a small figure. Preferably,in this class of compressor, the fluid flow passes through the speed ofsound, and hence the normal shock position is located, in the statorblading to provide optimum performance. However, the compressor isoperable even though the fluid flow is diffused through the speed ofsound downstream of the stator blades after elficient high velocityfluid turning in the stator blades. The second, as a result of thedevelopment of the above described blading, is an axial flow compressorhaving a supersonic rotor wherein the normal shock is positioned in therotor and a subsonic relative fluid flow velocity is provided at theexit of the rotor blades, and having a supersonic stator, likewisepreferably having the normal shock position therein. Both the rotor andstator have a high degree of flow turn providing unusually high overallpressure ratio. I

It is therefore an object of the present invention to provide asupersonic compressor having an unusually high overall pressure ratio,in the range of 4 or more per stage.

Another object of the present invention is to provide a supersoniccompressor providing a high angle of fluid turn of fluid having avelocity which is at some point in the system above the speed of sound.

Still a further object of the present invention is to provide eflicientcompressor blading capable of difiusmg supersonic air flow through thespeed of sound and simultaneously turning the flow through a large anglewithprovide a novel impulse type supersonic compressor whereinsupersonic fluid flow velocities are imposed on the stator blading inletand the flow through the stator is efliciently turned to the axialposition and diffused through the speed of sound.

Yet another object of the present invention is to provide a highpressure ratio supersonic compressor providing a shock-in rotor andshock-in-stator flow pattern.

Still other and further features and objects of the present inventionwill at once become apparent to those skilled in the art from aconsideration of the attached drawings wherein two embodiments of thepresent invention are shown by way of illustration only, and wherein:

Figure 1 is a partial elevational view in cross-section through acompressor constructed according to the present invention;

Figure 2 is a fragmentary developed plan view of the rotor and statorblading of a first form of the present invention utilizing an impulserotor; and

Figure 3 is a fragmentary developed plan view of the rotor and statorblading of a second form of the invention utilizing a shock-in-rotor andshock-in-stator blading construction.

As shown in the drawings:

In the embodiment of the invention illustrated, a rotor is driven by arotating shaft 11 mounted in a housing 12 having a generally streamlinedopening annulus 13. Blades 14 are provided radially projecting from theperipheral surface of the rotor 14) and blades 15 are rigidly secured inthe housing 12 to provide a stator. The annular passage 16 leads fromthe trailing edge of the blades 15 to the outlet of the compressor. Itwill, of course, be understood that while these general components areknown in the art, and as thus broadly described, are

- shown in the prior copending application, the specific blading whichwill hereinafter be described, taken in combination with the specificconfiguration of the inner and outer annulus walls 16a and 1615,respectively, comprise that portion for which novelty is claimed.

As described above, the major problem fa ing the industry in developmentof highly eflicient compressors is the problem of efficient turning ofthe fluid at and above supersonic velocities. A high rate of fluidturning is extremely desirable in order to provide a high pressure ratiofor a single stage and in the absence of such turning several stages arenecessary in order to provide a pressure ratio on the. order of 5.Accordingly, while the rotor and stator blading illustrated anddescribed in the above mentioned copending application provide a veryhigh efficiency as a result of its novel construction accuratelycontrolling fluid flow contraction and positioning of the normal shockin the rotor passages, it provides a relatively lower pressure ratio perstage of approximately 2.5. In the operation of the compressor set forthin the above identified copending application, it was found that anincrease of fluid flow turn substantially beyond approximately 12 causedsevere fluid flow separation .at the exit end of the rotor blading. Thisflow separation and turbulence quickly dropped the overall efliciency ofcompression to a figure well below desirable limits and accordingly theproblem of high fluid turn without such flow separation become theproblem from which the application stems.

4 The present blading has, exclusive of the impulse rotor, leading edgecharacteristics substantially identical to those set forth in thecopending application, and hence is provided with a small blade includedangle of approximately 4 and has provision for several increases inblade wedge angle with substantially the maximum channel contractionthat will accept fluid flow. This provides supersonic turning withseveral oblique shock Waves upstream of a normal shock wave positionedat or slightly downstream of the blade passage throat. As a result ofactual tests made on large numbers of passage shapes and contours thecritical control factor in providing efli cient high velocity fluid turnin supersonic blading was discovered to be the rate of area expansionprovided downstream of the point of maximum channel contraction, orsubsequent to the normal shock position.

Contrary to the understanding of those heretofore working in the art,effective high turn may be accomplished by limiting the area expansionof the fluid flow channel to a very small amount. For situations whereinan inlet velocity to the rotor blading is Mach 1.6 or above it has beenfound that an area expansion rate of less than 3% will preventseparation of the fluid flow and hence provide eflicient operation. Thisfigure of 3% may be very slightly increased with a decrease in the inletvelocity, V below Mach 1.6 without substantially affecting in an adverseway the efliciency of the blading.

This discovery, when applied to compressors provided with an inletvelocity relative to the blading, V above the speed of sound or aboveMach 1.0 permits high velocity fluid turn efliciently in two alternativemanners. The first of these is high rate of turning without sub stantialdecrease in velocity and accordingly substantially no change in fluidflow passage from the inlet to the outlet of the rotor blading. Such acompressor would provide a rotor operating as an impulse rotor and thevelocity of the entering fluid would never drop to the speed of sound orbelow during its passage through the rotor. By placing the blades fairlyclose together to provide relatively high solidity and by providing asmoothly increasing wedge angle on the high pressure side of each bladeefiicient supersonic fluid turn has been accomplished without providinga plurality of strong oblique shock waves. An illustration of thisarrangement is found in Figure 2 wherein the blades 14 of the rotor 10are provided with sharp leading edges 14a followed by a curved surfaceforming a continually increasing wedge angle. No throat or constrictionis provided in the rotor blading shown in Figure 2 and in fact, veryslight area expansion may be provided in order to compensate forboundary layer build-up adjacent the surfaces of the blades.Accordingly, fluid entering the compressor along the axial lineindicated by the vector V will have a relative velocity relative to theblading of V at entry into the blades and will have an exit velocity Vrelative to the blades which is substantially as great as V and whichwill be supersonic. Considering the vector diagram illustrating thevelocities of the fluids entering the stator blading 15 it will be notedthat the vectorial sum of the velocities V and V,, the latter indicatingthe velocity of the blades occasioned by rotation of the rotor, providesa relative velocity entering the stator V which is extremely high. Inpractice, when a rotor of the impulse type, above described, is utilizedfor providing a flow turn in the rotor of approximately 100 with a rotorblade velocity of 1400 feet per second and an inlet relative velocity Vof approximately Mach pulse type rotor blading above described, theextremely high velocities entering the stator increase the problem ofturning the fluid flow to the axial direction and at the same timediffusing the fluid. Using the principles above discussed, it has beenfound that efficient stator blading may be provided, capable ofsatisfactorily turning the fluid flow from its vectorial direction Vshown in Figure 2 to the axial direction with simultaneous diffusionthrough the speed of sound. This is accomplished through the provisionof stator blades having a sharp leading edge and a general supersonicflow contraction pattern completely set forth and described in the abovementioned copending application. Thus, a gradually increasing wedgeangle is provided on the high pressure side of the stator blades 15,providing a plurality of oblique shock waves which diffuse the fluidflow, with some simultaneous turning, down to the speed of sound atwhich point the normal shock occurs. Subsequent to the normal shock thefluid flow is turned through a large angle to the axial directionsubsonically with further diffusion accomplished through the provisionof an expansion of the flow area downstream of the normal shock. Thisexpansion is, as above described, limited to approximately 3% and whenso limited it has been found that the fluid flow does not separate.

It will be understood that because of the oblique shocks developedduring supersonic contraction of the fluid flow, through the use of aseries of sharp increases in blade wedge angle, more eflicient diffusionof the velocity of the fluid from supersonic to the subsonic is providedthan would be provided where a single normal shock wave is developed.However, it is noted that tests and operation of compressors constructedaccording to the present invention show that the best total pressurerecovery, or overall efliciency, of supersonic blading of the type hereunder discussion, results from operation at the highest static pressureratio for the blades. Accordingly, positioning of the normal shock atthe earliest possible point in the flow channel, with subsequentexpansion and diffusion provides optimum results. This condition issomewhat in conflict with the use of a plurality of oblique shock wavessince the construction of the blading to provide such waves ordinarilyrequires more blade length prior to the positioning of the normal shockthan is necessary if the normal shock is almost immediately set up.Accordingly, it is preferred that the oblique shock waves if utilized,be set up relatively close to each other in order to move the normalshock forward as far as possible.

If space limitations should prevent the use of blading sufliciently longto provide a substantial pressure rise through diifusion, with thenominal expansion areas permitted, and it accordingly becomes necessaryto choose between the efliciencies effected through the use of obliqueshock wave patterns and the efiiciency eiiected through moving thenormal shock forward in the flow channel it has been found that theeffect of moving the normal shock forward is superior and that viscosityand secondary flow losses are decreased more than enough for movement ofthe normal shock wave forward to overbalance the additional loss throughelimination of the oblique shocks. Accordingly, in such an instance arapid flow contraction adjacent the inlet to the blades, setting up thenormal shock almost immediately followmg the bow wave, such as forexample at 20 in Figure 2 is preferred. The alternative situation, inwhich oblique shocks are provided is satisfactory, however, and is shownin dotted lines with the normal shock 20a lying downstream of one ormore oblique shocks 2%, all lying behind the initial shock wave 21.

In Figure 3 both the rotor and stator operate transonically, in otherwords, the blades of both the rotor and the stator operate to reduce therelative velocity of the fluid flowing therethrough from a supersonicrelative inlet velocity to a subsonic relative outlet velocity. in orderto provide a sufficiently high subsonic velocity relative to the rotorbladings at their outlet to cause a supersonic inlet velocity to thestator, it is necessary that a high angle of turning be provided in therotor.- This angle of turning need not, however, be as great in theembodiment shown in Figure 3 as in the impulse rotor arrangement shownin Figure 2 since a pressure rise occurs in the rotor as well as in thestator and accordingly a lower inlet velocity at the stator blades willprovide the same overall compressor rotor and stator combination,pressure ratio. However, in View of the fact that the greater the turnprovided, the greater the total pressure ratio will be, it is desiredthat a high degree of fluid flow turn be provided. Accordingly, theblading shown in Figure 3 is set up to provide a fluid flow turn ofapproximately in the rotor, which is shown as the angle A between thevector quantities V and V The rotor blades provide initial flowcontraction, with flow turn through a plurality of oblique shock waves14b, and subsequent expansion limited to approximately 3% as abovedescribed. The stator is constructed to diffuse and turn the fluid flowentering the stator blading, namely V to the axial direction andaccordingly, as shown, the stator blading provides a fluid turn ofapproximately 60 indicated by the angle B. As in the case of the rotor,the stator blades provide flow contraction through oblique shock waves20c and normal shock wave 23, with subsequent 3% flow area expansion.The turning provided by both rotor and stator is, of course,substantially greater than heretofore considered practicable insupersonic operation and as described above has been achieved throughcontrol of the expansion of the channel area downstream of the normalshock condition.

As may be seen from consideration of the vector diagrams in Figure 3,although both the vector quantities V and V which comprise the actualinlet fluid flow velocity to the rotor and the relative outlet velocityfrom the rotor may be subsonic, the velocity component added by rotationof the rotor itself in the usual supersonic operating ranges of from1200 to 1500 f.p.s. will provide a relative fluid velocity V to therotor blading well above the speed of sound and a velocity V to thestator blading likewise well above the speed of sound. As a result oftest operations it has been found that a compressor operating with botha transonic rotor and transonic stator with controlled expansion areadownstream of the normal shock location, and with the normal shocklocation positioned in the early portion of the flow channel, asindicated at 22 and 23, respectively, an overall pressure ratio of above4, with an overall pressure recovery or efliciency of 70% or better isaccomplished.

The diagrams of Figures 2 and 3 do not, of course,

illustrate structural features incorporated in the compressor forinsuring radial equilibrium. It is, of course, essential for eflicientoperation of any compressor that the fluid flow have radial equilibriumin order to prevent excessive pressure gradients radially along the flowpassages. This equilibrium may be provided as shown in Figure l byproviding radial outward turn of the compressor flow passages. As willbe noted from a consideration of Figure 1, the inner radial surface ofthe channel 16a moves radially outwardly at a greater rate than theouter surfaces 16b. This difierence provides an apparent contractiontoward the downstream portion of the blading while passing through theblades 15. This contraction in radial dimension, however, isincorporated to limit the expansion that would otherwise be caused bythe necessary tapered construction of the trailing edges of thecompressor blades.

Downstream of the stator, the fluid, which is traveling in the axialdirection subsonically, is further diffused in an expanding channelhaving, as shown, an expansion rate of about 7%. 0

Additional radial equilibrium is, of course, obtained through a twistingof the individual rotor blades about their radial axis to compensate forthe increase of actual blade velocity with increasing radius and tothereby provide for substantially constant eflective inlet velocity Vthroughout the radial length of the rotor blades. In practice this isaccomplished through the provision of a decreasing angle of blade attackwith increase in rotor radius, the angle of attack being the anglebetween the low pressure side of the individual blade adjacent itsleading edge and the fluid relative inlet flow. This is demonstrated atC in Figure 2.

It will thus be apparent to those skilled in the art that we haveprovided a novel supersonic compressor capable of providing overallpressure ratio of approximately 4 to 1 in a single rotor-stator stage.This extremely high pressure ratio, heretofore considered impossible insupersonic compressor construction, has not only been achieved but hasbeen accomplished while maintaining a total pressure recovery orefficiency of at least 70% thereby providing an extremely practicalcompressor.

It will be understood, of course, that variations and modifications maybe made in the structures hereinbefore set forth by way of illustrationand accordingly it is intended that the scope of the present inventionbe limited solely by that of the appended claims.

We claim as our invention:

1. In combination in a supersonic compressor, a rotor having aperipheral surface spaced from an annular internally facing casingsurface by a plurality of radially extending rotor blades and a statorcomprising an annular passage coaxial with said rotor and having aplurality of radially extending stator blades therein, first meansdirecting inlet fluid to said rotor blading at a velocity relativethereto above Mach 1, said rotor blades having a turn in excess ofapproximatley along their axial length and a degree of fluid channelarea expansion in the downstream direction less than 3%, whereby fluidentering said rotor leaves said rotor traveling at a velocity relativeto said stator above a velocity of Mach 1 and having a vectorialmovement in the same direction as the rotation of said rotor, saidstator having a blade inlet relative velocity above Mach 1 and saidstator blades likewise having a turn in excess of approximately 10 alongtheir axial lengths whereby the fluid leaving said rotor blades isturned to the axial direction by flow between adjacent blades, secondmeans associated with said stator for difiusing the fluid flaw betweenthe blades thereof to a velocity of Mach 1 and third means associatedwith said stator for subsequently simultaneously substantially turningand dilfusing said flow subsonically without introducing flowseparation.

2. In combination in a supersonic compressor, a rotor having aperipheral surface spaced from an annular internally facing casingsurface by a plurality of radially extending rotor blades and a statorcomprising an annular passage coaxial with said rotor and having aplurality of radially extending stator blades therein, first meansdirecting inlet fluid to said rotor blading at a velocity relativethereto above Mach 1, said rotor blades having a turn in excess ofapproximately 10 along their axial length and a degree of fluid channelarea expansion in the downstream direction less than 3%, whereby fluidentering said rotor leaves said rotor traveling at a velocity relativeto said stator above a velocity of Mach 1 and having a vectorialmovement in the same direction as the rotation of said rotor, saidstator having a blade inlet relative velocity above Mach 1 and saidstator blades likewise having a turn in excess of approximately 10 alongtheir axial lengths whereby the fluid leaving said rotor blades isturned to the axial direction by flow between adjacent blades, secondmeans associated with said stator for difiusing the fluid flow betweenthe blades thereof to a velocity of Mach 1 and third means associatedwith said stator for. subsequently simultaneously substantially turningand diffusing said flow subsonically without introducing flowseparation, said last named means including surfaces on adjacent statorblades and said stator inner and outer annular walls providing across-sectional area expansion in the downstream direction from a pointof minimum area to the outlet of said stator blade passageways on theorder of 3%.

3. In combination in a supersonic compressor, a rotor havinga peripheralsurface spaced from an annular internally facing casing surface by aplurality of radially extending rotor blades and a stator comprising anannular passage coaxial with said rotor and having a plurality ofradially extending stator blades therein, first means directing inletfluid to said rotor blading at a velocity relative thereto above Mach 1,said rotor blades having a turn in excess of approximately 10 alongtheir axial length and providing an initial contraction of said fluidflow with resultant diflusion to a velocity of Mach 1 and subsequentfluid flow expansion on the order of 3% whereby fluid entering saidrotor leaves said rotor traveling at a velocity relative to said statorabove a velocity of Mach 1 and having a movement in the same directionas the rotation of said rotor, said stator blades likewise having a turnin excess of approximately 10 along their axial length whereby the fluidleaving said rotor blades is turned to the axial direction by flowbetween adjacent stator blades, second means associated with said statorfor contracting the fluid flow between the blades thereof for diffusionof the fluid to a velocity of Mach 1 and third means associated withsaid stator for subsequently simultaneously turning said fluid to theaxial direction and expanding said flow subsonically.

4. In combination in a supersonic compressor, a rotor having aperipheral surface spaced from an annular internally facing casingsurface by a plurality of radially extending rotor blades and a statorcomprising an annular passage coaxial with said rotor and having aplurality of radially extending stator blades therein, first meansdirecting inlet fluid to said rotor blading at a velocity relativethereto above Mach 1, said rotor blades having a turn in excess ofapproximately 10 along their axial length and providing an initialcontraction of said fluid flow with resultant diflusion to a velocity ofMach 1 and subsequent fluid flow expansion on the order of 3% wherebyfluid entering said rotor leaves said rotor traveling at a velocityrelative to said stator above a velocity of Mach 1 and having a movementin the same direction as the rotation of said rotor, said stator bladeslikewise having a turn in excess of approximately 10 along their axiallength whereby the fluid leaving said rotor blades is turned to theaxial direction by flow between adjacent stator blades, second meansassociated with said stator for contracting the fluid flow between theblades thereof for diflusion of the fluid to a velocity of Mach 1 andthird means associated with said stator for subsequently simultaneouslysubstantially turning and expanding said flow subsonically, said lastnamed means including surfaces on adjacent stator blades and said statorinner and outer annular walls providing a cross-sectional area expansionin the downstream direction from a point of maximum contraction to theoutlet of said stator on the order of 3%.

5. In a supersonic compressor a plurality of blades forming incombination with inner and outer annular walls a plurality of fluid flowchannels, means introducing fluid to said channels at a velocityrelative to said blades above Mach 1, means on adjacent blades forcontracting said fluid to thereby difius'e said fluid to a velocity ofMach 1 and means on said blades downstream of the point of maximumcontraction for simultaneously turning and expanding said fluidsubsonically through a turn in excess of approximately 10, said meanslimiting said expansion to 3%.

6.- In a supersonic compressor, a plurality of blades forming incombination with a pair of annular wall surfaces a plurality of fluidflow channels, means introducing fluid to said channels at a velocityabove Mach 1 relative thereto, means on adjacent blades for contractingsaid fluid flow and thereby diffusing said flow to a velocity of Mach 1and means associated with said blades and said walls downstream of thepoint of maximum fluid flow contraction for simultaneously causing fluidflow turn in excess of approximately 10 and an expansion thereof on theorder of 3%.

7. In a supersonic compressor, a pair of blades forming in combinationwith a pair of annular wall surfaces, a flow channel, means introducingfluid to said channel at a relative inlet velocity above Mach 1, meanson the opposing faces of said blades for contracting said fluid flow atthe maximum rate said flow may be contracted without blocking said fiowto thereby position the normal shock resulting from said contraction ata point adjacent 10 the leading edges of said blades, and means on saidopposed blade faces and said walls downstream of the point of maximumfluid flow contraction for simultaneously turning said fluid flowthrough an angle in excess of approximately 10 and expanding said flowat a rate on the order of 3%.

References Cited in the file of this patent UNITED STATES PATENTS2,435,236 Redding Feb. 3, 1948 2,623,688 Davidson Dec. 30, 19522,628,768 Kantrowitz Feb. 17, 1953 FOREIGN PATENTS 687,365 Great BritainFeb. 11, 1953

